The present invention relates generally to gas turbine engines, and, more specifically, to turbine flowpaths therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the hot gases in turbine stages for powering the compressor and producing output work in driving an upstream fan in a turbofan aircraft engine application, or in powering an external drive shaft for marine and industrial applications.
A high pressure turbine (HPT) immediately follows the combustor outlet and includes a stationary turbine nozzle followed in turn by a row of turbine rotor blades extending radially outwardly from the perimeter of a supporting disk. The HPT may include more than one stage and drives the rotor supporting the compressor rotor blades.
A low pressure turbine (LPT) follows the HPT and typically includes several stages of turbine nozzles and cooperating rows of rotor blades. The LPT drives the fan, or the external drive shaft.
Each of the turbine nozzles includes a row of stator vanes having corresponding airfoil configurations including generally concave pressure sides and opposite, generally convex suction sides extending axially in chord between opposite leading and trailing edges. The vane airfoils also extend in radial span between radially inner and outer bands or flow confining platforms which define the individual flow passages between adjacent vanes.
Similarly, each turbine rotor stage includes a row of turbine blades having integral dovetails mounted in complementary dovetail slots in the perimeter of the supporting rotor disk. Each rotor blade includes an airfoil portion extending radially outwardly from the dovetail at an intervening flow-bounding platform disposed at the root of the airfoil.
Each row of turbine blades is surrounded by a corresponding annular turbine shroud which provides a small clearance or gap above the radially outer tips of the blades to confine the combustion gases to flow in the individual flowpaths between adjacent turbine airfoils.
Since the first stage nozzle of the HPT first receives the hottest combustion gases from the combustor it is subject to substantial thermal expansion and contraction during operation. In order to reduce thermally induced stresses during operation, the turbine nozzle is circumferentially segmented into singlets of individual vanes and corresponding band segments, or doublets of two vanes supported in corresponding band segments.
The nozzle segments are suitably mounted together in an annular ring and have corresponding endfaces which define the circumferential joints therebetween in the form of axially extending splitlines. Each endface includes an axial slot therein which receives a spline seal for sealing the splitlines. In this way, the segmented annular nozzle is free to expand and contract during operation at the corresponding axial splitlines for minimizing thermally induced stress therein.
Similarly, each turbine rotor blade is individual mounted in a corresponding dovetail in the perimeter of the supporting rotor disk and is free to expand and contract without restraint from adjacent turbine rotor blades for minimizing stresses due to thermal expansion and contraction during operation. Since the turbine shrouds are suspended around the blade tips, relative expansion and contraction therebetween is permitted which correspondingly affects the size of the tip clearance.
The turbine nozzles and rotor blades are specifically designed in aerodynamic contour and flowpath dimensions for maximizing performance of the gas turbine engine. In the individual flowpaths between the nozzle vanes, a throat of minimum flow area is created with the flowpath converging to the throat and then diverging aft therefrom towards the corresponding row of turbine blades. Similarly, each flowpath defined between adjacent turbine blades also has a throat of minimum flow area, with the flowpath converging to the throat and then diverging in the downstream aft direction.
The inter-vane flowpaths are bound by the radially inner and outer integral bands. And the inter-blade flowpaths are bound by the integral blade platforms at the root of the airfoil and the radially outer turbine shroud.
Accordingly, the design of a modern gas turbine engine addresses in substantial detail the aerodynamic contours of the individual nozzle vanes and turbine blades from root to tip, as well as the corresponding aerodynamic contours of the inner and outer bands in the nozzle, and the platforms and turbine shrouds in the rotor stage. Modern three-dimensional computational techniques are now available for minutely analyzing the aerodynamic contours of all these components for maximizing engine performance.
However, the aerodynamic contours of the turbine nozzles and rotor stages are nevertheless subject to the small manufacturing tolerances of the individual components, and the small assembly tolerances when the components are mounted in the engine. These tolerances necessarily introduce random differences in the relative positions of adjacent components due to the independent assembly and mounting of these components. These random differences in final position correspondingly affect aerodynamic performance of the turbine, and in turn the overall efficiency of the engine.
In particular, the inboard surfaces of the inner and outer nozzle bands are precisely configured for maximizing aerodynamic performance of the nozzle, but the axial splitlines between adjacent nozzle segments affect nozzle performance. Since the nozzle vanes have typical airfoil configurations, they turn or bend the combustion gas flow during their axial downstream passage between the inter-vane flowpaths.
Accordingly, the combustion gases will firstly traverse the axial splitlines in one direction near the leading edges of the vanes and then traverse the same splitlines in an opposite direction near the trailing edges as the gas streamlines bend around the vanes during operation.
The flowpath surfaces of the inner and outer bands are manufactured for a nominally flush assembled position to ensure smooth flow of the combustion gas streamlines over the axial splitlines during operation. However, since the nozzle segments are subject to manufacturing and assembly tolerances, corresponding radial steps will be formed at the splitlines which will vary randomly in magnitude and direction within the small tolerances.
An up-step between nozzle segments will correspondingly obstruct combustion gas flow and reduce performance as well as locally heat the protruding step. A down-step is more desirable since it is subject to less heating from the combustion gases than an up-step, but nevertheless introduces a surface discontinuity which decreases efficiency.
Furthermore, since the combustion gas streamlines change direction as they flow over an individual splitline between the leading and trailing edges of the vanes, a desirable down-step at either the leading or trailing edges of the vanes will necessarily result in an undesirable up-step at the opposite side of the segment for nominally flush band designs.
Similarly, the turbine rotor blades have integral platforms which define the inner flowpath surface between the blades, which platforms are subject to similar steps between platforms subject to manufacturing and assembly tolerances.
The prior art includes various solutions for the inter-airfoil step problem in turbine nozzles and rotor stages. These solutions include providing chamfers on the platform or band edges, and varying the radial position of those edges to ensure primarily only down-steps as the flow streamlines bend in their downstream travel between the leading and trailing edges of the airfoils.
However, the use of chamfers and height changes in the platforms or bands correspondingly affects aerodynamic performance of the stator and rotor stages, and a compromise in design must be made for minimizing overall losses.
In particular, modern three-dimensional computational analysis now permits detailed design of the flowpath surfaces of the nozzle bands and blade platforms for maximizing efficiency of the inter-airfoil flowpaths. Changes in those flowpath surfaces should be eliminated or minimized in addressing the inter-airfoil stepping problem for maximizing efficiency of the engine.
Accordingly, it is desired to provide an inter-airfoil configuration in a gas turbine engine for resolving the steps between bands or platforms while minimizing aerodynamic contour changes of the bands and platforms themselves.